Compact, low pressure-drop shock-driven combustor and rocket booster, pulse detonation based supersonic propulsion system employing the same

ABSTRACT

A supersonic propulsion system is provided. The supersonic propulsion system includes a plurality of systems for efficiently creating cyclic detonations and at least one rocket booster device. Each of the systems include at least a first initiator chamber configured to generate an initial wave, at least one main chamber coupled to the first initiator chamber. The main chamber is configured to generate a main wave and to output products of supersonic combustion. The products are generated within the main chamber. The main chamber is configured to enable the main wave to travel upstream and downstream within the main chamber when the first initiator chamber is located outside the main chamber. The system further includes an initial connection section located between the first initiator chamber and the main chamber that enhances a combustion process via shock focusing and shock reflection.

PRIORITY

The present application is a continuation-in-part application of U.S.application Ser. No. 11/346,714, filed Feb. 6, 2006, the entiredisclosure of which is incorporated herein by reference.

BACKGROUND OF THE INVENTION

This invention relates generally to cyclic pulsed detonation combustors(PDCs) and more particularly, to a rocket booster, pulse detonationbased supersonic propulsion system employing compact, low pressure-dropshock driven combustors.

In a generalized pulse detonation combustor, fuel and oxidizer (e.g.,oxygen-containing gas such as air) are admitted to an elongatedcombustion chamber at an upstream inlet end of the pulse detonationcombustor. An igniter (spark or plasma ignitor) is used to initiate acombustion process within the pulse detonation combustor. Following asuccessful transition to detonation, a detonation wave propagates towardan outlet of the pulse detonation combustor at supersonic speed causinga substantial combustion of the fuel and oxidizer mixture before themixture can be substantially driven from the outlet. A result of thecombustion is to rapidly elevate pressure within the pulse detonationcombustor before a substantial amount of gas can escape through theoutlet. An effect of this inertial confinement is to produce nearconstant volume combustion. The pulse detonation combustor can be usedto produce pure thrust or can be integrated in a gas-turbine engine. Theformer is generally termed a pure thrust-producing device and the latteris generally a hybrid engine device. A pure thrust-producing device isoften used in a subsonic or supersonic propulsion vehicle system, suchas, rockets, missiles, and an afterburner of a turbojet engine.Industrial gas turbines are often used to provide output power to drivean electrical generator or motor. Other types of gas turbines may beused as aircraft engines, on-site and supplemental power generators, andfor other applications.

A deflagration-to-detonation transition (DDT) process begins when amixture of fuel and air in the chamber is ignited via a spark, laser orother source. A subsonic flame kernel generated from the ignitionaccelerates as the subsonic flame travels along the length of thechamber due to chemical processes and flow mechanics. As the subsonicflame reaches critical supersonic speeds, “hot spots” are created thatcreate localized explosions, eventually transitioning the subsonic flameto a super-sonic detonation wave. The DDT process can take up to severalmeters of the length of the chamber, and efforts have been made toreduce the distance used for DDT by using internal obstacles, such asorifice plates or spirals, in the flow of a mixture of fuel and oxidizerwithin the chamber. However, the obstacles for cyclic detonation deviceshave a relatively high pressure drop and are cooled. Moreover, thedetonation initiation, in the chamber with obstacles, occurs within arun-up length which ranges from and including 15 to 20 times a diameterof the chamber, and thus the run-up length increases with increasingchamber diameter. For practical propulsion systems, the run-up lengthdue to this constraint can be prohibitively long.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, a system for efficiently creating cyclic detonations isprovided. The system includes at least a first initiator chamberconfigured to generate an initial wave, at least one main chambercoupled to the first initiator chamber. The main chamber is configuredto generate a main wave and to output products of supersonic combustion.The products are generated within the main chamber. The main chamber isconfigured to enable the main wave to travel upstream and downstreamwithin the main chamber when the first initiator chamber is locatedoutside the main chamber. The system further includes an initialconnection section located between the first initiator chamber and themain chamber that enhances a combustion process via shock focusing andshock reflection.

In another aspect, a system for generating power and/or thrust isdescribed. The system includes an oxidizer supply system comprising acompressor configured to compress an oxidizer, a fuel supply systemcomprising a pump configured to pressurize fuel, at least a firstinitiator coupled to the oxidizer supply and the fuel supply system, andconfigured to generate an initial wave. The system further includes amain chamber coupled to the first initiator chamber. The main chamber isconfigured to generate a main wave, and configured to receive oxidizerfrom the compressor and fuel from the pump, where the main chamber isconfigured to output power generated from the initial wave generatedwithin the first initiator chamber. The main chamber is configured toenable the main wave to travel upstream and downstream within the mainchamber when the first initiator chamber is located outside the mainchamber.

In yet another aspect, a method for generating power/thrust isdescribed. The method includes coupling a main chamber to a firstinitiator chamber, generating an initial wave within the first initiatorchamber, generating a main wave within the main chamber, configuring themain wave to travel upstream and downstream within the main chamber uponconfiguring the first initiator chamber to be located outside the mainchamber, and outputting from the main chamber thrust generated from theinitial wave.

In an additional aspect of the present invention, a plurality of theabove described systems for creating cyclic detonations are used alongwith a rocket booster device in a supersonic propulsion system foraviation applications. For this application, the normal combustion orpropulsion core of a supersonic propulsion system is replaced with aplurality of the above described systems and a rocket booster device,such that a supersonic propulsion system is provided which can operateat cruise speeds of Mach 2-5, at altitudes ranging from 50,000-80,000feet.

As used herein, a “pulse detonation combustor” PDC (also including PDEs)is understood to mean any device or system that produces both a pressurerise and velocity increase from a series of repeating detonations orquasi-detonations within the device. A “quasi-detonation” is asupersonic turbulent combustion process that produces a pressure riseand velocity increase higher than the pressure rise and velocityincrease produced by a deflagration wave. Embodiments of PDCs (and PDEs)include a means of igniting a fuel/oxidizer mixture, for example afuel/air mixture, and a detonation chamber, in which pressure wavefronts initiated by the ignition process coalesce to produce adetonation wave. Each detonation or quasi-detonation is initiated eitherby external ignition, such as spark discharge or laser pulse, or by gasdynamic processes, such as shock focusing, auto ignition or by anotherdetonation (i.e. cross-fire).

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a block diagram of an exemplary embodiment of a system forefficiently creating cyclic detonations.

FIG. 2 shows a cross-section of an embodiment of a main chamber and aninitial initiator chamber included within the system of FIG. 1.

FIG. 3 is a block diagram of another exemplary embodiment of a systemfor efficiently creating cyclic detonations.

FIG. 4 shows a cross-section of the main chamber, the initial initiatorchamber, and an additional initiator chamber included within the systemof FIG. 3.

FIG. 5 shows an isometric view of the main chamber, the initialinitiator chamber, and the additional initiator chamber.

FIG. 6 shows an isometric view of a system including a plurality ofconnection sections included within the system of FIG. 3.

FIG. 7 is a schematic of an exemplary gas turbine engine including atleast one of the systems of FIGS. 1 and 3.

FIG. 8 is a schematic diagram of an embodiment of a system for injectiona fuel and oxidizer mixture within the main chamber.

FIG. 9 illustrates an embodiment of a strut for injecting fuel withinthe main chamber.

FIG. 10 is a diagrammatical representation of an embodiment of asupersonic propulsion system in accordance with the present invention.

FIG. 11 is a cross-sectional representation of an embodiment of thesupersonic propulsion system shown in FIG. 10.

FIG. 12 is a cross-sectional representation of an additional embodimentof a supersonic propulsion system of the present invention.

FIG. 13 is a diagrammatical representation of an exit portion of anembodiment of a main chamber of the present invention.

FIG. 14 is a cross-sectional representation of a further embodiment of asupersonic propulsion system of the present invention, which includes arocket booster device.

FIG. 15 is a diagrammatical representation of an exit portion of theembodiment shown in FIG. 14.

DETAILED DESCRIPTION OF THE INVENTION

A shock-to-detonation Transition (SDT) can be used to initiatedetonations in a large combustion chamber by transitioning a supersonicflame or a detonation wave, which is generated in an ignitor with asmall diameter, into a larger main chamber filled with fuel-oxidizermixture. In the SDT process, one or more initiators generate asupersonic flame or a detonation wave using a deflagration-to-detonationtransition (DDT) process, which is then transitioned to a larger chamberusing a carefully positioned reflecting and shock-focusing surface. Theresulting supersonic flame or detonation wave propagates through thelarger main chamber consuming a fuel-air mixture within the larger mainchamber. In general, the SDT process makes detonation transition from asmaller chamber to a larger chamber possible, and a run-up time and arun-up length used for this transition are much smaller when compared tothe DDT process occurring in a chamber.

A pulse detonation combustor (PDC) includes a device or system thatproduces pressure rise, temperature rise and velocity increase from aseries of repeating detonations or quasi-detonations within the device.A quasi-detonation includes a supersonic turbulent combustion processthat produces pressure rise, temperature rise and velocity increasehigher than pressure rise, temperature rise and velocity increaseproduced by a deflagration wave. Embodiments of PDCs include a fuelinjection system, an oxidizer flow system, a means of igniting afuel/oxidizer mixture, and a detonation chamber, in which pressure wavefronts initiated by the ignition process coalesce to produce adetonation wave. Each detonation or quasi-detonation is initiated eitherby external ignition, such as spark discharge or laser pulse, or by gasdynamic processes, such as shock focusing, autoignition or by anotherdetonation (cross-fire). The geometry of the PDC is such that thepressure rise of the detonation wave expels combustion products, such ascombustion gases, out the pulse detonation combustor exhaust to producea thrust force. Pulse detonation combustion can be accomplished in anumber of types of PDCs including shock tubes, resonating detonationcavities and tubular/tuboannular/annular combustors. As used herein, theterm “chamber” includes pipes having circular or non-circularcross-sections and having constant or varying cross sections. Exemplarychambers include cylindrical tubes, as well as tubes having polygonalcross-sections, for example hexagonal tubes.

FIG. 1 is a block diagram of an exemplary embodiment of a system 100 forefficiently creating cyclic detonations. Examples that use cyclicdetonations include systems to produce a propulsive force and thrust.System 100 includes an initial initiator chamber 102, a main chamber104, a plurality of fuel supplies 106 and 108, a plurality of oxidizersupplies 110 and 112, a plurality of fuel injectors 114 and 116, aplurality of valves 118 and 120, a controller 122, a plurality ofcontroller output lines 124, 126, and 128, a plurality of fuel supplylines 130 and 132, a plurality of oxidizer supply lines 134 and 136, andan initial initiation device 138. A cross-section of main chamber 104and initial initiator chamber 102 along a line A-A is shown in FIG. 2.As an example, a length to diameter ratio of main chamber 104 rangesfrom 0.5 to 5. Initial initiator chamber 102 is located outside mainchamber 104.

Referring back to FIG. 1, main chamber 104 includes a hollow gap 140 andinitial chamber includes a hollow gap 142. As used herein, the term“controller” is not limited to just those integrated circuits referredto in the art as a controller, but broadly refers to a processor, amicroprocessor, a microcontroller, a programmable logic controller, anapplication specific integrated circuit, and another programmablecircuit.

Each of fuel supplies 106 and 108 may be a tank that stores fuel, suchas a liquid fuel, such as, but not limited to, gasoline, diesel fuel,butane, pentane, hexane, jet fuel (JP 10), or Jet-A fuel. In analternative embodiment, fuel supply 106 stores gaseous fuel, such asethylene or hydrogen. In one embodiment, each oxidizer supply 110 and112 is an air tank that stores air. In an alternative embodiment,oxidizer supply 110 and 112 can include air from atmosphere or caninclude exhaust air from an engine, such as a turbine engine. Examplesof each of fuel injectors 114 and 116 include, but are not limited tobeing, an effervescent atomizer, a flash vaporizing injector, apressure-assist atomizer, an air-assist atomizer, and a supercriticalliquid injector. Each of valves 118 and 120 includes a solenoid valve.Initial initiation device 138 can be, but is not limited to being, aspark plug, a plasma igniter, and/or a laser source. In the exemplaryembodiment, each controller output line 124, 126, and 128 is aconducting medium, such as a metal wire.

Main chamber 104 extends from a point 144 to a point 146, extends from apoint 148 to a point 150, and extends from point 144 to point 148.Initial initiator chamber 102 extends from a point 152 to a point 154,extends from a point 156 to a point 158, and extends from point 152 topoint 156. Points 144 and 148 are closer to a head-end 157 of mainchamber 104 than points 146 and 150. Initial initiator chamber 102 iscoupled to main chamber 104 via an initial transition or initialconnection section 160 that extends from point 154 to a point 162 andfrom point 158 to point 164. Main chamber 104 does not includeobstacles, such as, an orifice plate, a spiral, a portion of initialinitiator chamber 102, and a portion of initial connection section 160.In one embodiment, a length of main chamber 104 between points 150 and162 is longer than a length of main chamber 104 between points 148 and164. In another embodiment, a length of main chamber 104 between points150 and 162 is the same as a length of main chamber 104 between points148 and 164. The length of main chamber 104 is measured substantiallyparallel to a z-axis.

Initial connection section 160 is integrated with main chamber 104. Forexample, initial connection section 160 is attached to and locatedoutside main chamber 104. Each of initial initiator chamber 102, mainchamber 104, and initial connection section 160 are fabricated from ametal, such as stainless steel or aluminum. Main chamber 104 is parallelto initial initiator chamber 102. Alternatively, main chamber 104 is notparallel to initial initiator chamber 102. For example, main chamber 104forms an angle ranging from zero degrees to 179 degrees within initialinitiator chamber 102. In another alternative embodiment, initialinitiator chamber 102 may have a spiral shape or has a zigzag shape.Main chamber 104 is coupled, such as bolted or welded, to initialconnection section 160 and initial connection section 160 is coupled,such as bolted or welded, to initial initiator chamber 102.

The shape of the initial connection section 160 is configured to enhanceshock reflection and shock focusing to enhance initiation in the mainchamber 104. Initial connection section 160 includes a sharp edge 166forming an angle 168 ranging from five degrees to and including 90degrees between initial initiator chamber 102 and a side 167 of initialconnection section 160. Alternatively, initial connection section 160includes a curved edge instead of or in addition to sharp edge 166. Inanother alternative embodiment, initial connection section 160 includesa recessed cone or a paraboloid instead of or in addition to sharp edge166. In another alternative embodiment, initial connection section 160includes more than one, such as between 2 and 5, sharp edges. Initialconnection section 160 is a piece other than main chamber 104 andinitial initiator chamber 102.

Controller 122 sends an “on” signal via controller output line 124 tovalve 118 and an “on” signal via controller output line 126 to valve120. Upon receiving an “on” signal from controller 122, valve 118actuates or opens. Similarly, upon receiving an “on” signal viacontroller output line 124, valve 120 actuates or opens. When valve 118is open, fuel stored within fuel supply 106 is supplied via fuel supplyline 130 to fuel injector 114. Fuel injector 114 atomizes fuel receivedvia fuel supply line 130 into a plurality of droplets and supplies thedroplets to initial initiator chamber 102. Alternatively, if fuelinjector 114 is not included within system 100, fuel, such as liquid orgaseous fuel, is supplied from fuel supply line 130 to initial initiatorchamber 102. Fuel is supplied from fuel supply 106 to initial initiatorchamber 102 in a pulsed manner at a specific frequency or alternativelyis continuously supplied to initial initiator chamber 102 for a specificperiod of time. Additionally, a flow of oxidizer is supplied fromoxidizer supply 110 via oxidizer supply line 134 to initial initiatorchamber 102.

Fuel is supplied from fuel supply 108 to main chamber 104 continuouslyfor a period of time. Additionally, a flow of oxidizer is supplied fromoxidizer supply 112 via oxidizer supply line 136 to main chamber 104.When valve 120 is open, fuel from fuel supply 108 is supplied via fuelsupply line 132 to fuel injector 116, where the fuel is atomized into aplurality of droplets, which are channeled into main chamber 104.

After determining that a pre-determined amount of time has passed sincevalves 118 and 120 were opened, controller 122 transmits an “off” signalto valve 118 via controller output line 124 and an “off” signal to valve120 via controller output line 126. Valves 118 and 120 close uponreceiving “off” signals. Supply of fuel to main chamber 104 and toinitial initiator chamber 102 stops upon closure of valves 118 and 120.In an alternative embodiment controller 122 does not transmit an “off”signal to close valve 120. In the alternative embodiment, valve 120remains open during each cycle of generation of a wave within initialinitiation chamber 102.

Controller 122 sends a signal to initial initiation device 138 viacontroller output line 128. Upon receiving the signal via controlleroutput line 128, initial initiation device 138 creates a spark withininitial initiator chamber 102. The spark within initial initiatorchamber 102 ignites a mixture of fuel and oxidizer within initialinitiator chamber 102 to generate an ignition kernel. The timing of thespark created within initial initiation chamber 102 can be before orafter the valves 118 and 120 receive ‘off’ signals from controller 122.The ignition kernel within initial initiator chamber 102 expands into adeflagration flame that accelerates into a turbulent flame and aninitial wave, such as a shock wave, a quasi-detonation wave, or adetonation wave, within initial initiator chamber 102. The initial wavewithin initial initiator chamber 102 propagates through a mixture offuel and oxidizer within initial initiator chamber 102 to increase thepressure within initial initiator chamber 102.

The initial wave generated within initial initiator chamber 102 travelsfrom initial initiator chamber 102 via initial connection section 160 tomain chamber 104. Initial connection section 160 may have a reflectivesurface that is designed to focus and/or reflect the initial wavetowards main chamber 102. The oxidizer and fuel mixture within mainchamber 104 is ignited by the initial wave generated within initialinitiator chamber 102. The oxidizer and fuel mixture within main chamber104 is ignited to generate a main wave within main chamber 104. The mainwave has a near sonic velocity, such as ranging from Mach 0.8 to Mach 1,or a supersonic velocity. The main wave travels upstream towards an end157 of main chamber 102 and also travels downstream towards an end 170of main chamber 102. End 157 is located opposite to end 170. The mainwave generated within main chamber 104 propagates through a mixture offuel and oxidizer within main chamber 104 to increase the pressurewithin main chamber 104. The combustion gases formed within main chamber104 exit main chamber 104 via end 170 of main chamber 104 to generatethrust and or power. End 170 is open to enable the combustion gasesgenerated within main chamber 104 to exit main chamber 102 via end 170and end 169 is closed to prevent the combustion gases from exiting mainchamber 102 via end 169.

FIG. 3 is a block diagram of another exemplary embodiment of a system300 for generating power/thrust. System 300 includes initial initiatorchamber 102, main chamber 104, fuel supplies 106 and 108, a fuel supply302, oxidizer supplies 110 and 112, an oxidizer supply 304, fuelinjectors 114 and 116, a fuel injector 306, valves 118 and 120 and avalve 308, controller 122, controller output lines 124, 126, and 128, aplurality of controller output lines 310 and 312, fuel supply lines 130and 132, a fuel supply line 314, oxidizer supply lines 134 and 136, anoxidizer supply line 316, initial initiation device 138, an additionalinitiation device 318, and an additional initiator chamber 320.Additional initiator chamber 320 is located outside main chamber 104. Inone embodiment, main chamber 104 has the same diameter as at least oneof initial initiator chamber 102 and additional initiator chamber 320.In an alternative embodiment, main chamber 104 has a larger diameterthan at least one of initial initiator chamber 102 and additionalinitiator chamber 320. In another alternative embodiment, main chamber104 has a smaller diameter than at least one of initial initiatorchamber 102 and additional initiator chamber 320. Fuel supply 302 may bea tank that stores fuel, such as the liquid fuel. In an alternativeembodiment, fuel supply 302 stores gaseous fuel. In an alternativeembodiment, system 300 may not include fuel injector 306. Additionalinitiator chamber 320 includes a hollow gap 322. A cross-section of mainchamber 104, initial initiator chamber 102, and additional initiatorchamber 320 along a line B-B is shown in FIG. 4. Moreover, an isometricview of main chamber 104, initial initiator chamber 102, and additionalinitiator chamber 320 is shown in FIG. 5.

Referring back to FIG. 3, oxidizer supply 304 is an oxidizer tank thatstores oxidizer. In an alternative embodiment, oxidizer supply 304 caninclude air from atmosphere or can include exhaust air from an engine,such as a turbine engine. Fuel injector 306 includes, but is not limitedto being, an effervescent atomizer, a flash vaporizing injector, apressure-assist atomizer, an air-assist atomizer, and a supercriticalliquid injector. Valve 308 includes a solenoid valve. Additionalinitiation device 318 can be, but is not limited to being, a spark plug,a plasma igniter, and/or a laser source. In the exemplary embodiment,each controller output line 310 and 312 is a conducting medium, such asa metal wire.

Additional initiator chamber 320 extends from a point 324 to a point326, extends from a point 328 to a point 330, and extends from point 324to point 328. Additional initiator chamber 320 is coupled to mainchamber 104 via an transition section or additional connection section332 that extends from point 326 to a point 334 and from point 330 to apoint 336. The shape of the additional connection section 332 isconfigured to enhance shock reflection and shock focusing to enhanceinitiation in the main chamber 104. Additional connection section 332 isintegrated with main chamber 104. For example, additional connectionsection 332 is attached to and located outside main chamber 104. Mainchamber 104 does not include other obstacles, such as, a portion ofadditional initiator chamber 320 and a portion of additional connectionsection 332. Additional connection section 332 is a piece other thanmain chamber 104 and additional initiator chamber 320. Additionalinitiator chamber 320 and additional connection section 332 arefabricated from a metal, such as stainless steel or aluminum. Mainchamber 104 is parallel to additional initiator chamber 320.Alternatively, main chamber 104 is not parallel to additional initiatorchamber 320. For example, main chamber 104 forms an angle ranging fromzero degrees to 179 degrees within additional initiator chamber 320. Inanother alternative embodiment, additional initiator chamber 320 mayhave a spiral shape or has a zigzag shape. Main chamber 104 is coupled,such as bolted or welded, to additional connection section 332 andadditional connection section 332 is coupled, such as bolted or welded,to additional initiator chamber 320.

Additional connection section 332 includes a sharp edge 338 forming anangle 340 ranging from five degrees to and including 90 degrees betweenadditional initiator chamber 320 and a side 342 of additional connectionsection 332. Alternatively, additional connection section 332 includes acurved edge instead of or in addition to sharp edge 338. In anotheralternative embodiment, additional connection section 332 includes arecessed cone instead of or in addition to sharp edge 338. In yetanother alternative embodiment, additional connection section 332includes a paraboloid instead of or in addition to sharp edge 338. Inanother alternative embodiment, additional connection section 332includes more than one, such as between 2 and 5, sharp edges.

Controller 122 sends an “on” signal via controller output line 310 tovalve 308. Upon receiving an “on” signal from controller 122, valve 308actuates or opens. When valve 308 is open, fuel stored within fuelsupply 302 is supplied via fuel supply line 314 to fuel injector 306.Fuel injector 306 atomizes fuel received via fuel supply line 314 into aplurality of droplets and supplies the droplets to additional initiatorchamber 320. Alternatively, if fuel injector 306 is not included withinsystem 300, fuel, such as liquid or gaseous fuel, is supplied from fuelsupply line 314 to additional initiator chamber 320. Fuel is suppliedfrom fuel supply 302 to additional initiator chamber 320 in a pulsedmanner at a certain frequency or alternatively is continuously suppliedto additional initiator chamber 320 for a period of time. Additionally,a flow of oxidizer is supplied from oxidizer supply 304 via oxidizersupply line 316 to additional initiator chamber 320.

After determining that an amount of time has passed since valve 308 wasopened, controller 122 transmits an “off” signal to valve 308 viacontroller output line 310. Valve 308 closes upon receiving an “off”signal. Supply of fuel to additional initiator chamber 320 stops uponclosure of valve 308. In an alternative embodiment, controller 122 doesnot transmit an “off” signal to close valve 308. In the alternativeembodiment, valve 308 remains open during each cycle of generation of awave within additional initiation chamber 320.

Controller 122 sends a signal to additional initiation device 318 viacontroller output line 312. Upon receiving the signal via controlleroutput line 312, additional initiation device 318 creates a spark withinadditional initiator chamber 320. In one embodiment, the spark withinadditional initiator chamber 320 is created at the same time at whichthe spark within initial initiator chamber 102 is created. For example,controller 122 sends a signal to initial initiation device 138 at thesame time at which controller 122 sends a signal to additionalinitiation device 318. In an alternative embodiment, the spark withininitial initiator chamber 102 is generated at a different time than atime at which the spark within additional initiator chamber 320 isgenerated. As an example, the sparks within initial initiator chamber102 and additional initiator chamber 320 are alternated by creating thespark within initial initiator chamber 102 at a first time, creating aspark within additional initiator chamber 320 at a second time followingthe first time, and creating a spark within initial initiator chamber102 at a third time following the second time. As another example, thesparks within initial initiator chamber 102 and additional initiatorchamber 320 are alternated by creating the spark within initialinitiator chamber 102 at the first time, creating a spark withinadditional initiator chamber 320 at the second time consecutive to thefirst time, and creating a spark within initial initiator chamber 102 atthe third time consecutive to the second time. As yet another example,the spark within initial initiator chamber 102 is repeated at a firstperiod and the spark within additional initiator chamber 320 is repeatedat a second period, where a time of repetition of the first period doesnot coincide with a time of repetition of the second period. As yetanother example, controller 122 sends a signal to initial initiationdevice 138 at a different time than a time of sending a signal toadditional initiation device 318. The timing of the spark created withinadditional initiation chamber 320 can be before or after the valves 308and 120 receive ‘off’ signals from controller 122.

The spark within additional initiator chamber 320 ignites a mixture offuel and oxidizer within additional initiator chamber 320 to generate anignition kernel. The ignition kernel within additional initiator chamber320 expands into a deflagration flame that accelerates into a turbulentflame and an additional wave, such as a shock wave, a quasi-detonationwave, or a detonation wave, within additional initiator chamber 320. Theadditional wave generated within additional initiator chamber 320propagates through a mixture of fuel and oxidizer within additionalinitiator chamber 320 to increase the pressure within additionalinitiator chamber 320.

The additional wave generated within additional initiator chamber 320travels from additional initiator chamber 320 via additional connectionsection 332 to main chamber 104. Additional connection section 332 isdesigned to focus and/or reflect the additional wave towards mainchamber 102. The oxidizer and fuel mixture within main chamber 104 isignited by the additional wave generated within additional initiatorchamber 320 and the initial wave generated within initial initiatorchamber 102. Alternatively, the liquid fuel received from fuel injector306 is transformed into a gaseous form within main chamber 104 by theinitial wave generated within initial initiator chamber 102, and theoxidizer and fuel mixture within main chamber 104 is ignited by theadditional wave generated within additional initiator chamber 320. Inanother alternative embodiment, the liquid fuel received from fuelinjector 306 is transformed into a gaseous form by the initial wavegenerated within initial initiator chamber 10 and is conditioned fordetonation by the initial wave. The oxidizer and fuel mixture withinmain chamber 104 is ignited by the initial and additional waves frominitial initiator chamber 102 and additional initiator chamber 320 togenerate the main wave within main chamber 104, and the combustion gasesformed within main chamber 104 exit main chamber 104 via end 170 of mainchamber 104 to generate thrust and or power.

It is noted that system 300 may include more than one additionalinitiator chamber 320. For example, system 300 includes three additionalinitiator chambers and initial initiator chamber 102.

FIG. 6 shows an isometric view of a system 600 including a plurality ofconnection sections 602 and 604. Connection section 602 is an example ofinitial connection section 160 and connection section 604 is an exampleof additional connection section 332. System 600 includes a plurality ofports 606, 608, and 610. Initial initiator chamber 102 is attached, suchas welded or bolted, to port 606 and additional initiator chamber 320 isattached, such as welded or bolted, to port 610. Main chamber 104 isattached, such as welded or bolted, to port 608. Main chamber 104 iscoupled to initial initiator chamber 102 via connection section 602 andis coupled to additional initiator chamber 320 via connection section604.

It is noted that any of main chamber 104, initial initiator chamber 102,and additional initiator chamber 320 may include a liner and a coolant.The coolant flows between the liner and a side wall of any of mainchamber 104, initial initiator chamber 102, and additional initiatorchamber 320.

FIG. 7 is a schematic of an exemplary gas turbine engine 700 including alow pressure compressor 702, a high pressure compressor 704, and apressure-rised combustion system 706. Engine 700 also includes a highpressure turbine 708 and a low pressure turbine 710. Low pressurecompressor 702 and low pressure turbine 710 are coupled by a first shaft712, and high pressure compressor 704 and high pressure turbine 708 arecoupled by a second shaft 714. In one embodiment, engine 700 is aF110-129 engine available from General Electric Aircraft Engines,Cincinnati, Ohio. Pressure-rised combustion system 706 includes at leastone system 100 except that the at least one system 100 is controlled bycontroller 122. Alternatively, pressure-rised combustion system 706includes at least one system 300 except that the at least one system 300is controlled by controller 122.

In operation, air flows through low pressure compressor 702 from aninlet side 716 of engine 700 and is supplied from low pressurecompressor 702 to high pressure compressor 704 to generate compressedair. Compressed air is delivered to any of oxidizer supply lines 134,136, and oxidizer supply line 316. In an alternative embodiment, air issupplied to any of oxidizer supply lines 134, 136, and oxidizer supplyline 316 from low pressure compressor 702. In another alternativeembodiment, air is supplied to any of oxidizer supply lines 134, 136,and oxidizer supply line 316 from low pressure turbine 710. In yetanother alternative embodiment, air is supplied to any of oxidizersupply lines 134, 136, and oxidizer supply line 316 from a combinationof low pressure compressor 702 and low pressure turbine 710. Similarly,in still another alternative embodiment, air is supplied to any ofoxidizer supply lines 134, 136, and oxidizer supply line 316 from atleast one of high pressure compressor 704 and high pressure turbine 708.

Compressed air is mixed with fuel, such as fuel pressurized by a pump,and ignited to generate the combustion gases. The combustion gasesgenerated within pressure-rised combustion system 706 are channeled frompressure-rised combustion system 706 to drive turbines 708 and 710 andprovide thrust from an outlet 718 of engine 700. In an alternativeembodiment, any of systems 100, 300, and 704 can be, but are not limitedto being, used for other propulsion applications, such as, rocketboosters, rocket engines, missiles, and an unmanned combat aerialvehicle (UCAV).

FIG. 8 is a schematic diagram of an embodiment of a system 800 forinjecting a fuel and oxidizer mixture within main chamber 104. Aplurality of struts 802, 804, 806, 808, 810, and 812 are arrangedradially along a circumference of main chamber 104. Struts 802, 804,806, 808, 810, and 812 are arranged radially at head end 157 of mainchamber 104. Struts 802, 804, 806, 808, 810, and 812 are parallel to anxy plane formed by an x-axis and a y-axis. Alternatively, struts 802,804, 806, 808, 810, and 812 may not be parallel to the xy plane. Aplurality of angles between struts 802, 804, 806, 808, 810, and 812 areequal. As an example, an angle 814 between struts 810 and 812 is equalto an angle 816 between struts 802 and 812. In an alternativeembodiment, the plurality of angles between struts 802, 804, 806, 808,810, and 812 are unequal. For example, angle 814 is unequal to angle816. In another alternative embodiment, any number of struts, such as,2, 3, 4, 8, 9, or 10, can be arranged radially to inject a mixture offuel and oxidizer into main chamber 104.

Struts 802, 804, 806, 808, 810, and 812 are radially arranged tofacilitate a radial injection of a mixture of fuel and oxidizer intomain chamber 104. A mixture of fuel and oxidizer is radially injectedinto main chamber 104 via an opening or port 818 of strut 802, anopening or port 820 of strut 804, an opening or port 822 of strut 806,an opening or port 824 of strut 808, an opening or port 826 of strut810, and an opening or port 828 of strut 812. The radial injection offuel within main chamber 104 can be synchronous with an injection offuel into one of initial initiation chamber 102 and additionalinitiation chamber 320. The radial injection of fuel generates a finelyatomized spray resulting in a uniform mixture of fuel and oxidizerwithin main chamber 104. Moreover, the radial injection concentratesfuel at a point 830, such as a center, of main chamber 104. However, inthe present invention the radial injection is not limited to occurringat center 830 and can also occur at any point from the main chamber wallto at or near the center 830. The radial injection can be continuous oralternatively periodic, such as pulsed.

FIG. 9 illustrates an embodiment of a strut 902, which is an example ofone of struts 802, 804, 806, 808, 810, and 812. Strut 902 has an airfoilshape that minimizes a drag experienced by a mixture of fuel andoxidizer flowing within strut 902.

Technical effects of the herein described systems and methods forgenerating power include foregoing using an initiation device todirectly ignite the mixture of fuel and oxidizer within main chamber104. Other technical effects include using shock reflection and shockfocusing to raise the temperature and pressure within the connectionsection 160 and main chamber 104. This causes the mixture of fuel andoxidizer to ignite within a short initiation delay ranging from andincluding 0.1 millisecond (ms) to 5 ms, and propagate at supersonicspeeds. This short initiation delay is caused because of the hightemperature and pressure of the fuel-air mixture in the high amount ofinitiation energy. Further technical effects include assisting in abreak-up of a plurality of droplet of the liquid fuel within mainchamber 104 and conditioning the droplets for detonation. Still furthertechnical effects include achieving a repetitive or cyclical frequencyof generation of the main wave within main chamber 104. An example ofthe cyclical frequency includes a range between 0 to 100 Hertz. Othertechnical effects include providing a substantially unobstructed flow ofthe main wave within main chamber 104. Location of at least one ofinitial initiator chamber 102, additional initiator chamber 320, initialconnection section 160, and additional connection section 332 outsidemain chamber 104 facilitates the unobstructed flow.

Turning now to FIG. 10, an embodiment of a supersonic propulsion system900 of the present invention is shown. The propulsion system 900includes a plurality of the above described system for efficientlycreating cyclic detonations 100. Because the operation of each thesystems 100 have been discussed in detail, those discussion will not berepeated. As shown, a plurality of the systems 100 are encased in acasing structure 901 which is cylindrical in shape. However, the presentinvention is not limited to having a cylindrical shape, as other shapesmay be used.

At the upstream end of the plurality of systems 100 is an air flowcontrol structure 903 which is used to supply and control air flow toeach of the systems 100. The air flow control structure 903 may be of aconfiguration to have separate and discrete air flow valve structuresfor each of the systems 100. In this configuration, air flow to each ofthe systems 100 can be controlled separately so as to be able to alloweach of the systems 100 to operate independently to each other. In afurther embodiment of the present invention, the air flow controlstructure 903 is an integrated common air flow structure which suppliesair flow to each of the systems 100. Such a configuration may be simplerfor control and operation. It is noted that the present invention is notlimited to the structure and operation of the air flow control structure903. The air flow for each of the systems 100 is to be controlled toensure desired operational characteristics and proper operation, asdescribed previously. An upstream plenum 904 separates the inlet flowcontrol structure from the air inlets at the front of the outerframe/structure. This plenum serves the purpose of minimizing theflow/pressure fluctuations which propagate upstream, and also serves toseparate the unsteady downstream end of the PDE tubes from the steadyupstream inlet nozzles.

Air flow for the air flow control structure 903 may be obtained from airflow inlets (not shown) on the casing structure 901. The air flow inletsmay be positioned in the nose of the casing structure, or alternativelyalong the side portion of the casing structure 901.

At the end of the supersonic propulsion system 900 is a exit nozzleportion 905. In the embodiment shown in FIG. 10, the exit nozzle portion905 is of a straight nozzle type, which simply directs the exhaust fromthe systems 100 aft, through the exit nozzle portion 905. In anotherembodiment of the present invention, the exit nozzle portion 905 is avariable geometry nozzle. In such an embodiment, the geometry of thenozzle is controlled to optimize the performance of the supersonicsystem 900. Because the operation and structure of variable geometryexit nozzles are known, a detailed discussion will not be incorporatedherein. In another embodiment of the present invention, the exit nozzleportion is a converging diverging nozzle. The operation and structure ofconverging-diverging nozzles are well known and will not be discussed indetail herein. Because the exhaust gases of the plurality of systems 100is very high (above Mach 2) the converging-diverging nozzle is to bedesigned to optimize exhaust performance.

As shown in FIG. 10, the main chambers 104 of each of the systems 100contain an exit nozzle portion 907. As shown in FIG. 10, each of theexit nozzle portions 907 are converging-diverging nozzles. In anembodiment of the present invention the exit nozzle portions 907 mayalso be converging or diverging nozzles. Although shown in FIG. 10, eachof the exit nozzle portions 907 are the same for each system 100, it iscontemplated that at least some of the systems 100 have an exit nozzleportion 907 of one type and the other systems 100 have exit nozzleportions 907 of a different type.

In an embodiment of the invention, each of the exit nozzle portions 907of the systems 100 share a common exit plane. Thus, as shown in FIG. 10,the exit area of each of the nozzle portions 907 exist in the sameplane, which is perpendicular to the centerline of the propulsion system900. In another embodiment of the present invention, the exit nozzleportions 907 of the systems 100 are “staggered.” In this embodiment, theexit opening of at least some of the exit nozzle portions 907 aredifferent than the exit openings of others of the nozzle portions 907.

In the embodiment shown in FIG. 10, the main chambers 104 of the systems100 are approximately 40 inches in length. However, the presentinvention is not limited to this length, as other lengths may beutilized to optimize performance and operation. Additionally, in analternative embodiment, it is contemplated that the length of at leastsome of the main chambers 104 are different than that of others. Thevarying length of the main chambers 104 can provide the staggeredexhaust of the systems 100 into an exhaust plenum or the exit nozzleportion 905 of the casing 901.

FIG. 11 is a cross-sectional representation of the embodiment of theinvention shown in FIG. 10. In this embodiment, there are six systems100 distributed in a radially symmetric pattern. However, the presentinvention is not limited to this number of systems 100, nor the shownconfiguration. The number of systems 100 and their respectivedistribution can be optimized for performance and operationalcharacteristics. The distribution and configuration of the systems 100should be such so as to provide optimal packing efficiency.

In the present invention, packing efficiency (“η”) is the ratio of areaof the tubes over area of the plenum (or inner cavity of the casing901). In an embodiment of the invention the packing efficiency is over0.7. In the embodiment shown in FIG. 10, the packing efficiency is about0.72, including the center tube 909.

The packing efficiency affects the core mass flow losses of thepropulsion system 900. The higher the packing efficiency, the lower thelosses and the higher the system efficiency.

In an embodiment of the invention, the main chambers 104 have an innerdiameter of 6 inches and the inner diameter of the initiator chambers102 is 2 inches, providing a main chamber diameter to initiator chamberdiameter ratio of 3 to 1. In a further embodiment, the diameter of themain chambers 104 are up to 6.5 inches. As shown in FIG. 11, thediameter of all of the main chambers 104 are the same. In an alternativeembodiment, it is contemplated that one or more of the main chambers 104have a diameter which is different than the others. For example, it iscontemplated that half of the main chambers 104 have a diameter of 6inches, while the other half have diameters which are larger or smaller.It is noted that these are exemplary dimensions and the presentinvention is not limited to these specific exemplary sizes.

Each of the main chambers 104 is coupled to a initiator chamber 102, asdiscussed previously.

In FIG. 11, the center tube 909 is not a system 100, as described above.In an embodiment of the present invention, the center tube 909 is asupport structure providing structural support for the systems 100. In afurther embodiment, the tube 909 allows for cooling air/liquid to passthrough to assist in the cooling of the systems 100, as their operationgenerates a large amount of heat. In yet another alternative embodimentof the present invention, the center tube 909 is an additional system asdescribed above.

As shown in FIG. 11, each of the main chambers 104 is coupled to anindividual initiator chamber 102. Such a configuration allows theoperation and timing of the detonations in the chambers 104 to becontrolled. In one embodiment of the invention, the systems 100 areoperated in a synchronized fashion so that detonations occursimultaneously. In a further embodiment, the systems 100 are operatedsuch that the detonations in respective main chambers 104 occur atdifferent times. In one such embodiment, half of the main chambers 104are detonated at the same time, while the other half are in anotherstage of operation, such as purge, blowdown, fill, etc. In anotherembodiment, each of the main chambers 104 are detonated such that no onedetonation is at the same time. Stated differently, no two detonationsoccur at the same time. Such a configuration maximizes the operationalfrequency of the propulsion system 900, as the operational frequency ofthe system 900 is a product of the operational frequency of the systems100 and the number of systems 100.

Using the above described configuration, an embodiment of the presentinvention provides a propulsion system which uses the systems 100 thatcan achieve a cruise Mach number of at least 2.5 at altitudes of atleast 50,000 feet. A further embodiment of the present invention, canachieve a cruise Mach number in the range of 3 to 5, while at altitudesin the range of 50 kft to 80 kft.

In a further embodiment of the present invention, the operation of thevarious systems 100 are controlled based on the operational status ofthe propulsion system 900. For example, during a first stage ofoperation (climb, for example) all systems 100 are operating in asimultaneous detonation mode, while in a second stage of operation(cruise, for example) the systems 100 are operated out of phase witheach other, or at least some of the systems 100 are shut down, so only aportion are operating. The overall operation and control of the systems100 are such that the overall operation and performance goals of thepropulsion system 900 are achieved and optimized.

FIG. 12 is an additional embodiment of the present invention, which issimilar in structure to that shown in FIG. 11, except that the number ofinitiator chambers 102 has been reduced. In this embodiment, eachinitiator tube 102 supplies an initiation shock wave to at least two ofthe main chambers 104. Such a configuration aids in optimizing spacewithin the casing 901.

In one embodiment, the initiator chamber 102 supplies the initiationshock wave to each of its respective main chambers 104 at the same timeso as to detonation each of the main chambers 104 at the same time. Inan alternative embodiment, the flow from the initiator chambers 102 tothe main chambers 104 is valved or controlled so that the main chambers104 are operated out of phase with each other. Specifically, theinitiator flow is directed to a first one of the main chambers 104 totrigger detonation, while the other of the main chambers 104 is inanother portion of the cycle (such as blowdown, purge, fill, etc). Afterinitiating a detonation in the first main chamber 104, the initiatorchamber 102 is coupled to or valved to the next main chamber 1043 toinitiate detonation in that chamber. By alternating detonations betweenadjacent chambers (through the use of a single initiator chamber 102),cooling of the main chambers 104 becomes easier. In such an embodiment,the operational frequency of the initiator chamber 102 will need to behigher than the main chambers 104. In the embodiment described above,the frequency of the initiator chamber 102 is double that of the mainchambers 104.

The embodiment of the present invention shown in FIG. 12 may be operatedsimilarly to the embodiment shown in FIG. 11. Namely, they may beoperated in a synchronous or non-synchronous fashion. Further, theoperation of the respective systems 100 may be controlled based on theoperational status of the propulsion system 900.

It is noted that although FIG. 12 shows that each initiator chamber 102is coupled to two main chambers 104, the present invention is notlimited to this embodiment. It is contemplated that the initiatorchambers 102 may be coupled to more than two main chambers.

Further, it is also contemplated that any one main chamber 104 iscoupled to more than one initiator chambers 102.

FIG. 13 shows an embodiment of an exit nozzle portion 907 of a system100, at the downstream end of the main chamber 104. The exit nozzleportion 907 is a converging-diverging nozzle configuration. In a furtherembodiment, the nozzle 907 is either a converging or diverging nozzletype. In yet a further embodiment of the invention, it is contemplatedthat the exit nozzle portion 907 is of a straight nozzle type. Theembodiment shown in FIG. 13 is exemplary in nature, and it is noted thatthe present invention is not limited to this embodiment in any way.

As shown in FIG. 13, the exit nozzle portion 907 contains an exit plane911 through which the exhaust gases of the system 100 are expelled. Inan embodiment of the invention this exit plane 911 is upstream of theexit plane 913 of the casing 901 or exit nozzle portion 905. However, ina further embodiment of the invention, it is contemplated that at leastone of the main chambers 104 has an exit nozzle portion 907 whichextends beyond the exit plane 913 such that the exit plane 911 of thesystem 100 is downstream of the exit plane 911 of the casing 901 or exitnozzle portion 905.

FIG. 14 shows a cross-section of another embodiment of the presentinvention, where the supersonic propulsion system 900 further comprisesa rocket booster device 920.

Because of the operational characteristics of the systems 100 (describedabove), these systems 100 become most efficient at higher speeds (i.e.above Mach 2.5 or 3, typically). Thus, it is usually inefficient to usethe above described systems 100 at slower speeds. Therefore, it isdesirable to provide additional thrust to the propulsion system 900 inorder to assist the system 900 in reaching speeds at which the systems100 are more efficient.

The present invention accomplishes this by coupling the compact, lowpressure-drop shock driven combustor systems 100 with a rocket boosterdevice 920. The rocket booster device 920 is a solid fuel rocketpropellant booster device or a liquid fuel rocket propellant boosterdevice or a hybrid fuel rocket propellant booster device (e.g., liquidoxidizer and solid fuel). In an embodiment of the invention, the rocketbooster device is used to provide all or supplemental thrust throughall, or some, of launch, climb, transonic and top of climb portions of aflight envelope. In a further embodiment, the present invention is notlimited to using the rocket booster device 920 in only these portions ofthe flight envelope, as it is contemplated that the rocket boosterdevice can also be used in cruise and decent portions.

In an embodiment of the invention, the rocket booster device 920operates to propel the supersonic propulsion system 900 from Mach ˜1.0(air launching) to Mach 2.5 or above, at which time the systems 100 willtake over providing primary propulsion, up through cruise speeds of Mach3.0-5.0. It is contemplated that each of the rocket booster device 920and the systems 100 will overlap in their operation, so as to preventhaving a time when no propulsion is being provided. In a furtherembodiment of the present invention, the rocket booster device is of thetype which propels the propulsion system 900 from Mach 0 (surfacelaunch) to Mach 2.5 or above, at which time the systems 100 take over.

In the embodiment of the present invention shown in FIG. 14, the rocketbooster device 920 is positioned along the centerline of the propulsionsystem 900. In another embodiment, one or more rocket boosters devices920 are positioned radially outward from the systems 100, and may bepositioned external to the casing 901.

Further, as shown in FIG. 15, the rocket booster device 920 contains anexit nozzle 921 which directs the exhaust from the rocket booster device920 through the exit nozzle portion 905 of the propulsion system 900. Inan embodiment of the invention, an exit plane 923 of the rocket boosterexit nozzle 921 is co-planar with at least some of the exit planes 911of the nozzles 907. In a further embodiment, the exit plane of therocket booster device 920 exit nozzle is positioned either downstream orupstream of the exit planes 911 of the nozzles 907.

Further, in an additional embodiment, at least some of the systems 100,(shown radially in FIG. 14) are replaced with rocket booster devices920. For example, in the FIG. 14 embodiment, it is contemplated thatevery other device (shown radially around the centerline) is a rocketbooster device. Thus, in this embodiment, a combustion system 100 isadjacent to two rocket booster devices 920, and vice-versa.

In an embodiment with more than one rocket booster device 920, all ofthe rocket booster devices are operated simultaneously. However, in afurther embodiment, at least one rocket booster device 920 is operatedsuch that its operation (thrust time) ends either before or after theremaining rocket booster devices 920. This embodiment provides extendedimpulse time provide by the rocket booster devices 920.

As shown in FIG. 14, the cross-sectional dimensions of the rocketbooster device 920 is similar to that of the main chambers 104. However,the present invention is not limited in this regard. The overall sizeand geometry of the rocket booster device 920 is to be determined tosatisfy the desired operational and performance characteristics.

In an exemplary embodiment of the present invention, during theoperation of the rocket booster device 920 the systems 100 are notfunctioning. However, air flow is permitted to pass through at leastsome of the main chambers 104, so as to provide thermal cooling for therocket booster device 920. In this embodiment, the main chambers 104 androcket booster device 920 are thermally coupled so as to allow heattransfer.

While the invention has been described in terms of various specificembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theclaims.

1. A supersonic propulsion system; comprising: a casing structure; atleast one rocket booster device positioned either internally orexternally with respect to said casing structure; and a plurality ofsystems for creating cyclic detonations within said casing structure,wherein each of said systems comprise: at least a first initiatorchamber configured to generate an initial wave; at least one mainchamber coupled to said first initiator chamber, wherein said mainchamber is configured to generate a main wave and to output products ofsupersonic combustion, the products are generated within the mainchamber, wherein said main chamber is further configured to enable themain wave to travel upstream and downstream within said main chamber,and wherein said first initiator chamber is located outside said mainchamber; an initial connection section located between the firstinitiator chamber and the main chamber that enhances a combustionprocess via shock focusing and shock reflection.
 2. The propulsionsystem of claim 1, wherein said at least one rocket booster device is asolid fuel rocket propellant booster device, a liquid fuel rocketpropellant booster device, or a hybrid rocket fuel propellant boosterdevice.
 3. The propulsion system of claim 1, wherein said at least onerocket booster device is positioned along a centerline of said casingstructure.
 4. The propulsion system of claim 1, wherein a diameter ofsaid at least one rocket booster device is the same as at least some ofsaid main chambers.
 5. The propulsion system of claim 1, furthercomprising a plurality of said rocket booster devices, each of which ispositioned adjacent at least two of said systems.
 6. The propulsionsystem of claim 1, wherein during a first mode of operation said atleast one rocket booster device operates while none of said systemsoperate and during a second mode of operation at least one of saidsystems operates while said at least one rocket booster device does notoperate.
 7. The propulsion system of claim 1, wherein during a mode ofoperation said at least one rocket booster device and at least one ofsaid systems operate simultaneously.
 8. The propulsion system of claim1, wherein said rocket booster device comprises an exit nozzle having anexit plane, and at least one of said main chambers comprises an exitnozzle having an exit plane, and each of said exit planes are co-planar.9. A supersonic propulsion system, comprising: a casing structure; atleast one rocket booster device positioned either internally orexternally with respect to said casing structure; and a plurality ofsystems for generating thrust, each of said systems comprising: anoxidizer supply system comprising a compressor configured to compress anoxidizer; a fuel supply system comprising a pump configured topressurize fuel; at least a first initiator coupled to said oxidizersupply and said fuel supply system, and configured to generate aninitial wave; and a main chamber coupled to said first initiatorchamber, configured to generate a main wave, and further configured toreceive oxidizer from said compressor and fuel from said pump, whereinsaid main chamber is further configured to output power generated fromthe initial wave generated within said first initiator chamber, whereinsaid main chamber is further configured to enable the main wave totravel upstream and downstream within said main chamber, and whereinsaid first initiator chamber is located outside said main chamber. 10.The propulsion system of claim 9, wherein said at least one rocketbooster device is a solid fuel rocket propellant booster device, aliquid fuel rocket propellant booster device, or a hybrid rocket fuelpropellant booster device.
 11. The propulsion system of claim 9, whereinsaid at least one rocket booster device is positioned along a centerlineof said casing structure.
 12. The propulsion system of claim 9, whereina diameter of said at least one rocket booster device is the same as atleast some of said main chambers.
 13. The propulsion system of claim 9,further comprising a plurality of said rocket booster devices, each ofwhich is positioned adjacent at least two of said systems.
 14. Thepropulsion system of claim 9, wherein during a first mode of operationsaid at least one rocket booster device operates while none of saidsystems operate and during a second mode of operation at least one ofsaid systems operates while said at least one rocket booster device doesnot operate.
 15. The propulsion system of claim 9, wherein during a modeof operation said at least one rocket booster device and at least one ofsaid systems operate simultaneously.
 16. The propulsion system of claim9, wherein said rocket booster device comprises an exit nozzle having anexit plane, and at least one of said main chambers comprises an exitnozzle having an exit plane, and each of said exit planes are co-planar.